Gas turbine engine compressor arrangement

ABSTRACT

A gas turbine engine includes a low pressure compressor section and a high pressure compressor section. A low pressure turbine drives the low pressure compressor section. A gear arrangement is driven by the low pressure turbine to in turn drive a fan section. A pressure ratio across the low pressure compressor section is between about 4-8, and a pressure ratio across the high pressure compressor section is between about 8-15. In a separate feature, a compressor case includes a front compressor case portion and a rear compressor case portion, with the rear compressor case portion being axially further from an inlet case than the front compressor case portion. A support member extends between the fan section and the front compressor case portion.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.13/337,354, filed on Dec. 27, 2011, which is a continuation in-part ofU.S. patent application Ser. No. 13/294,492 filed on Nov. 11, 2011,which was a continuation of U.S. patent application Ser. No. 11/858,988filed on Sep. 21, 2007 now U.S. Pat. No. 8,075,261, and entitled “GasTurbine Engine Compressor Case Mounting Arrangement.”

BACKGROUND

The present invention relates generally to a compressor assembly in agas turbine engine.

Gas turbine engines are known, and typically include a compressor forcompressing air and delivering it downstream into a combustion section.A fan may move air to the compressor. The compressed air is mixed withfuel and combusted in the combustion section. The products of thiscombustion are then delivered downstream over turbine rotors, which aredriven to rotate and provide power to the engine.

The compressor includes rotors moving within a compressor case tocompress air. Maintaining close tolerances between the rotors and theinterior of the compressor case facilitates air compression.

Gas turbine engines may include an inlet case for guiding air into acompressor case. The inlet case is mounted adjacent the fan section.Movement of the fan section, such as during in-flight maneuvers, maymove the inlet case. Some prior gas turbine engine designs support afront portion of the compressor with the inlet case while anintermediate case structure supports a rear portion of the compressor.In such an arrangement, movement of the fan section may cause at leastthe front portion of the compressor to move relative to other portionsof the compressor.

Disadvantageously, relative movement between portions of the compressormay vary rotor tip and other clearances within the compressor, which candecrease the compression efficiency. Further, supporting the compressorwith the inlet case may complicate access to some plumbing connectionsnear the inlet case.

It would be desirable to reduce relative movement between portions ofthe compressor and to simplify accessing plumbing connection in a gasturbine engine.

Traditionally, a fan and low pressure compressor have been driven in oneof two manners. First, one type of known gas turbine engine utilizesthree turbine sections, with one driving a high pressure compressor, asecond turbine rotor driving the low pressure compressor, and a thirdturbine rotor driving the a fan. Another typical arrangement utilizes alow pressure turbine section to drive both the low pressure compressorand the fan.

Recently it has been proposed to incorporate a gear reduction to drivethe fan such that a low pressure turbine can drive both the low pressurecompressor and the fan, but at different speeds.

SUMMARY

In a featured embodiment, a gas turbine engine includes a fan sectionand a compressor section including both a low pressure compressorsection and a high pressure compressor section. A turbine sectionincludes a high pressure turbine driving the high pressure compressorsection and a low pressure turbine driving the low pressure compressorsection. A gear arrangement is driven by the low pressure turbine to inturn drive the fan section. An overall pressure ratio provided by thecombination of the low pressure compressor section and the high pressurecompressor section is provided by a pressure ratio across the lowpressure compressor section between about 4-8, and a pressure ratioacross the high pressure compressor section between about 8-15 toprovide the overall pressure ratio.

In another embodiment, the overall pressure ratio is above or equal toabout 35.

In another embodiment according to the foregoing embodiment, the overallpressure ratio is above or equal to about 40.

In another embodiment according to the foregoing embodiment, the overallpressure ratio is above or equal to about 50.

In another featured embodiment, a gas turbine engine comprises a fansection having a central axis and a compressor case for housing acompressor. An inlet case guides air to the compressor. The compressorcase is positioned axially further from the fan section than the inletcase. A support member extends between the fan section and thecompressor case with the support member restricting movement of thecompressor case relative to the inlet case. The compressor case includesa front compressor case portion and a rear compressor case portion. Therear compressor case portion is axially further from the inlet case thanthe front compressor case portion. The support member extends betweenthe fan section and the front compressor case portion, and the inletcase is removable from the gas turbofan engine separately from thecompressor case.

In another featured embodiment of the foregoing embodiment, thecompressor case includes a low pressure compressor section and a highpressure compressor section. A low pressure turbine drives the lowpressure compressor and a gear arrangement is driven by the low pressureturbine such that the gear arrangement drives the fan section. Anoverall pressure ratio provided by the combination of the low pressurecompressor section and the high pressure compressor section is above orequal to about 35.

In a further embodiment according to the foregoing embodiment, theoverall pressure ratio is above or equal to about 40.

In a further embodiment according to the foregoing embodiment, theoverall pressure ratio is above or equal to about 50.

In another embodiment according to the foregoing embodiment, thepressure ratio across the low pressure compressor section is betweenabout 4-8 and a pressure ratio across the high pressure compressorsection is between about 8-15.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic sectional view of a gas turbine engine.

FIG. 2 illustrates a sectional view of a prior art compressor casemounting arrangement. Notably, some aspects are not prior art.

FIG. 3 illustrates a sectional view of an example compressor casemounting arrangement of the current invention.

FIG. 4 illustrates a close up sectional view of the intersection betweenan inlet case and a low pressure compressor case.

FIG. 5 graphically shows a split in the compression ratios between thelow pressure and high pressure compressor sections.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 10including (in serial flow communication) a fan section 14, a lowpressure compressor 18, a high pressure compressor 22, a combustor 26, ahigh pressure turbine 30 and a low pressure turbine 34. The gas turbineengine 10 is circumferentially disposed about an engine centerline X.During operation, air is pulled into the gas turbine engine 10 by thefan section 14, pressurized by the compressors 18, 22 mixed with fuel,and burned in the combustor 26. Hot combustion gases generated withinthe combustor 26 flow through high and low pressure turbines 30, 34,which extract energy from the hot combustion gases.

In a two-spool design, the high pressure turbine 30 utilizes theextracted energy from the hot combustion gases to power the highpressure compressor 22 through a high speed shaft 38, and a low pressureturbine 34 utilizes the energy extracted from the hot combustion gasesto power the low pressure compressor 18 and the fan section 14 through alow speed shaft 42. However, the invention is not limited to thetwo-spool gas turbine architecture described and may be used with otherarchitectures such as a single-spool axial design, a three-spool axialdesign and other architectures. That is, there are various types of gasturbine engines, many of which could benefit from the examples disclosedherein, which are not limited to the design shown.

The example gas turbine engine 10 is in the form of a high bypass ratioturbine engine mounted within a nacelle or fan casing 46, whichsurrounds an engine casing 50 housing a core engine 54. A significantamount of air pressurized by the fan section 14 bypasses the core engine54 for the generation of propulsion thrust. The airflow entering the fansection 14 may bypass the core engine 54 via a fan bypass passage 58extending between the fan casing 46 and the engine casing 50 forreceiving and communicating a discharge airflow F1. The high bypass flowarrangement provides a significant amount of thrust for powering anaircraft.

The gas turbine engine 10 may include a geartrain 62 for controlling thespeed of the rotating fan section 14. The geartrain 62 can be any knowngear system, such as a planetary gear system with orbiting planet gears,a planetary system with non-orbiting planet gears or other type of gearsystem. The low speed shaft 42 may drive the geartrain 62. In thedisclosed example, the geartrain 62 has a constant gear ratio. It shouldbe understood, however, that the above parameters are only exemplary ofa contemplated geared gas turbine engine 10. That is, aspects of theinvention are applicable to traditional turbine engines as well as otherengine architectures.

The example engine casing 50 generally includes at least an inlet caseportion 64, a low pressure compressor case portion 66, and anintermediate case portion 76. The inlet case 64 guides air to the lowpressure compressor case 66.

As shown in FIG. 2, the low pressure compressor case 66 in an exampleprior art gas turbine engine 80 supports a plurality of compressorstator vanes 68. Notably, the low pressure compressor section 18, andthe high pressure compressor section 22, and the arrangement of therotors 70 and 170, respectively, are not part of the prior art. Aplurality of rotors 70 rotate about the central axis X, and, with thecompressor stator vanes 68, help compress air moving through the lowpressure compressor case 66. Downstream of the low pressure compressorthe air passes into the low pressure compressor at section 18 the airpasses into the high pressure compressor section 22, and is furthercompressed by its rotors 170. The mounting of the compressor as shown inFIG. 2 is prior art, however, the structure of the low pressurecompressor section 18 and high pressure compressor section 22, and therotors 70 and 170 were not part of the prior art.

A plurality of guide vanes 72 secure the intermediate case 67 to the fancasing 46. Formerly, the guide vanes 72 each included at least a rearattachment 74 and a forward attachment 78. The rear attachment 74connects to an intermediate case 76 while the forward attachment 78connects to the inlet case 64. The lower pressure compressor case 66 wasthus supported through the intermediate case 76 and the inlet case 64.

In the prior art, a plumbing connection area 82 is positioned betweenthe rear attachment 74 and the forward attachment 78. The plumbingconnection area 82 includes connections used for maintenance and repairof the gas turbine engine 80, such as compressed air attachments, oilattachments, etc. The forward attachment 78 extends to the inlet case 64from at least one of the guide vanes 72 and covers portions of theplumbing connection area 82. A fan stream splitter 86, a type of cover,typically attaches to the forward attachment 78 to shield the plumbingconnection area 82.

Referring now to an example of the present invention shown in FIG. 3, inthe turbine engine 90, the forward attachment 78 attaches to a frontportion of the low pressure compressor case 66. In this example, theforward attachment 78 extends from the guide vane 72 to support the lowpressure compressor case 66. Together, the forward attachment 78 andguide vane 72 act as a support member for the low pressure compressorcase 66. The plumbing connection area 82 (which includes connectionsused for maintenance and repair of the gas turbine engine 90, such ascompressed air attachments, oil attachments, etc) is positioned upstreamof the forward attachment 78 facilitating access to the plumbingconnection area 82. In contrast, the plumbing connection area of priorart embodiments was typically positioned between the rear attachment andthe forward attachment and the forward attachment typically extended tothe inlet case from at least one of the guide vanes, thereby coveringportions of the plumbing connection area, which complicated accessthereto; this complicated structure was further complicated by a fanstream splitter, a type of cover, that typically was attached to theforward attachment to shield the plumbing connection area.

In the embodiment shown in FIG. 3, an operator may directly access theplumbing connection area 82 after removing the fan stream splitter 86.The plumbing connection area 82 typically provides access to alubrication system 82 a, a compressed air system 82 b, or both. Thelubrication system 82 a and compressed air system 82 b are typically influid communication with the geartrain 62.

Maintenance and repair of the geartrain 62 may require removing thegeartrain 62 from the engine 90. Positioning the plumbing connectionarea 82 ahead of the forward attachment 78 simplifies maintenance andremoval of the geartrain 62 from other portions of the engine 90.Draining oil from the geartrain 62 prior to removal may take placethrough the plumbing connection area 82 for example. The plumbingconnection area 82 is typically removed with the geartrain 62. Thus, thearrangement may permit removing the geartrain 62 on wing or removing theinlet case 64 from the gas turbine engine 90 separately from the lowpressure compressor case 66. This reduces the amount of time needed toprepare an engine for continued revenue service, saving an operator bothtime and money.

Connecting the forward attachment 78 to the low pressure compressor case66 helps maintain the position of the rotor 70 relative to the interiorof the low pressure compressor case 66 during fan rotation, even if thefan section 14 moves. In this example, the intermediate case 76 supportsa rear portion of the low pressure compressor case 66 near a compressedair bleed valve 75.

As shown in FIG. 4, a seal 88, such as a “W” seal, may restrict fluidmovement between the inlet case 64 and the low pressure compressor case66. In this example, the seal 88 forms the general boundary between theinlet case 64 and the low pressure compressor case 66, while stillallowing some amount movement between the cases.

FIG. 5 shows a novel worksplit that has been invented to improve thefuel burn efficiency of a geared turbofan architecture with a fan 18connected to the low compressor 18 through a speed reduction device suchas a gearbox 62. Since a gear reduction 62 is incorporated between thefan 14 and the low pressure compressor 18, the speeds of the lowpressure compressor can be increased relative to a traditional two spooldirect drive arrangement. This provides freedom in splitting the amountof compression between the low pressure section 18 and the high pressuresection 22 that can be uniquely exploited to improve fuel burnefficiency on the geared turbofan architecture described in FIGS. 1 and2. This resulting worksplit is distinctly different from historical twoand three spool direct drive architectures as shown in FIG. 5.

Notably, while the gear train 62 is shown axially adjacent to the fan14, it could be located far downstream, and even aft of the low turbinesection 34. As is known, the gear illustrated at 62 in FIGS. 2 and 3could result in the fan 14 rotating in the same, or the oppositedirection of the compressor rotors 70 and 170.

It is known in prior art that an overall pressure ratio (when measuredat sea level and at a static, full-rated takeoff power) of at least 35:1is desirable, and that an overall pressure ratio of greater than about40:1 and even about 50:1 is more desirable. That is, after accountingfor the fan 18 pressure rise in front of the low pressure compressor 18,the pressure of the air entering the low compressor section 18 should becompressed as much or over 35 times by the time it reaches the outlet ofthe high compressor section 22. This pressure rise through the low andhigh compressors will be referred to as the gas generator pressureratio.

FIG. 5 shows the way that this high pressure ratio has been achieved inthe two prior art engine types versus the Applicant's engine'sconfiguration.

Area S₁ shows the typical operation of three spool arrangementsdiscussed the Background Section. The pressure ratio of the lowcompressor is above 8, and up to potentially 15. That is, if a pressureof 1 were to enter the low pressure compressor, it would be compressedbetween 8 to 15 times.

As can be further seen, the high pressure ratio compressor in thisarrangement need only compress a very low pressure ratio, and as low as5 to achieve a combined gas generator pressure ratio of above 35. Inaddition, the three spool design requires complex arrangements tosupport the three concentric spools.

Another prior art arrangement is shown at area S₂. Area S₂ depicts thetypical pressure ratio split in a typical two spool design with a directdrive fan. As can be seen, due to the connection of the fan directly tothe low pressure compressor, there is little freedom in the speed of thelow pressure compressor. Thus, the low pressure compressor can only do asmall amount of the overall compression. As shown, it is typically below4 times. On the other hand, the high pressure compressor must provide anamount of compression typically more than 20 times to reach the pressureratio of 40 (or 50).

The S₂ area results in undesirably high stress on the high pressurecompressor, which, in turn, yields challenges in the mounting of thehigh pressure spool. In other words, the direct drive system thatdefines the S₂ area presents an undesirable amount of stress, and anundesirable amount of engineering required to properly mount the highpressure spool to provide such high pressure ratios.

Applicant's current low compressor/high compressor pressure split isshow at area S₃. The fan is driven at a speed distinct from the lowpressure compressor, and a higher compression ratio can be achieved atthe low pressure compressor section than was the case at area S₂. Thus,as shown, the pressure ratio across the low pressure turbine may bebetween 4 and 8. This allows the amount of compression to be performedby the high pressure compressor to only need to be between 8 times and15 times.

The area S₃ is a enabling design feature that allows the geared turbofanarchitecture shown in FIGS. 1 and 2 to achieve very high gas generatorpressure ratio while avoiding the complexities of historical three spooland two spool direct drive architectures. The area S₃ is an improvementover both areas S₁ and S₂. As an example, a 3-4% fuel efficiency isachieved at area S₃ compared to area S₁. A fuel savings of 4-5% isachieved at area S₃, compared to area S₂.

In fact, in comparison to a gas turbine engine provided with a geardrive, but operating in the pressure ratios of area S₂, there is still a2% fuel burn savings at the S₃ area.

As such, the area S₃ reduces fuel burn, and provides engineeringsimplicity by more favorably distributing work between the hotter highpressure spools and colder low pressure spools.

Stated another way, the present invention provides a combination of alow pressure compressor and a high pressure compressor which togetherprovides an overall pressure ratio of greater than 35 and, in someembodiments, greater than 40, and even 50. This high overall pressureratio is accomplished by a beneficial combination of a pressure ratioacross the low pressure compressor of about 4-8, and a pressure ratioacross the high pressure ratio compressor of about 8-15.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a fan section; a compressor section,including both a low pressure compressor section and a high pressurecompressor section, said fan section delivering a portion of air into abypass path outwardly of the compressor section, and a portion of airinto the compressor section; a turbine section, including a highpressure turbine driving said high pressure compressor section, and alow pressure turbine driving said low pressure compressor section, and agear arrangement driven by said low pressure turbine to in turn drivethe fan section; and an overall pressure ratio provided by thecombination of said low pressure compressor section and said highpressure compressor section, with said overall pressure ratio beingprovided by a pressure ratio across said low pressure compressor sectionbetween about 4-8, and a pressure ratio across said high pressurecompressor section between about 8-15 to provide said overall pressureratio, said overall pressure ratio is above or equal to about
 50. 2. Anarrangement for a gas turbine engine comprising: a fan section having acentral axis; a compressor case for housing a compressor; an inlet casefor guiding air to said compressor, said compressor case positionedaxially further from said fan section than said inlet case; a supportmember extending between said fan section and said compressor casewherein said support member restricts movement of said compressor caserelative to said inlet case; and said compressor case includes a frontcompressor case portion and a rear compressor case portion, said rearcompressor case portion being axially further from said inlet case thansaid front compressor case portion, wherein said support member extendsbetween said fan section and said front compressor case portion, andsaid inlet case is removable from said gas turbofan engine separatelyfrom said compressor case, said compressor case includes a low pressurecompressor section and a high pressure compressor section, and wherein alow pressure turbine drives said low pressure compressor, and a geararrangement is driven by said low pressure turbine such that said geararrangement drives said fan section, and there being a plumbingconnection area to be utilized for maintenance and repair.
 3. Thearrangement as set forth in claim 2, wherein an overall pressure ratioprovided the combination of said low pressure compressor section andsaid high pressure compressor section being above or equal to about 35.4. The arrangement as set forth in claim 3, wherein the overall pressureratio is above or equal to about
 40. 5. The arrangement as set forth inclaim 4, wherein the overall pressure ratio is above or equal to about50.
 6. The arrangement as set forth in claim 3, wherein a pressure ratioacross said low pressure compressor section is between about 4-8, and apressure ratio across the high pressure compressor section is betweenabout 8-15.